Corresponding to the variety of purposes and goals possible for space ship flights, the demands placed on the vehicle will also be very different from mission to mission. For space ships, it will, therefore, be necessary to make the structure of the vehicle compatible with the uniqueness of the respective trip to a far greater extent than for the vehicles used for transportation to date. Nevertheless, the important equipment as well as the factors critical for the structure will be common for all space ships.
The external form of a space vehicle will have to be similar to that of a projectile. The form of a projectile is best suited for overcoming air drag at the high velocities attained by the vehicle within the Earth's atmosphere (projectile velocity, in accordance with previous statements!).
Fundamental for the internal structure of a rocket vehicle is the type of the propellants used. They must meet with the following requirements:
That they achieve an exhaust velocity as high as possible because the necessity was recognized previously for an expulsion velocity of the exhaust masses as high as possible.
That they have a density as high as possible (high specific weight), so that a small tank would suffice for storing the necessary amount of weight. Then, on the one hand, the weight of the tank is decreased and, on the other hand, the losses due to air drag also become smaller.
That their combustion be carried out in a safe way compatible with generating a constant forward thrust.
That handling them cause as few difficulties as possible.
Any type of gunpowder or a similar material (a solid propellant), such as used in fireworks rockets, would be the most obvious to use. The structure of the vehicle could then be relatively simple, similar to that of the familiar fireworks rocket. In this manner it would, no doubt, be possible to build equipment for various special tasks, and this would in particular pave the way for military technology, a point to be discussed below.
However for purposes of traveling in outer space, especially when the transportation of people is also to be made possible, using liquid propellants should offer far more prospects for development options, despite the fact that considerable engineering problems are associated with these types of propellants; this point will be discussed later.
The most important components of a space ship for liquid propellants are as follows: the propulsion system, the tanks for the propellants, the cabin and the means of landing. The propulsion system is the engine of the space ship. The reactive force is produced in it by converting the onboard energy stored in the propellant into forward thrust. To achieve this, it is necessary to pipe the propellants into an enclosed space in order to burn them there and then to let them discharge (exhaust) towards the rear. Two basic possibilities exist for this:
The same combustion pressure continuously exists in the combustion chamber. For the propellants to be injected, they must, therefore, be forced into the combustion chamber by overcoming this pressure. We will designate engines of this type as "constant pressure rocket engines."
The combustion proceeds in such a fashion that the combustion chamber is continuously reloaded in a rapid sequence with propellants, repeatedly ignited (detonated) and allowed to exhaust completely every time. In this case, injecting the propellants can also take place without an overpressure. Engines of this type we will designate as "detonation (or explosion) rocket engines."
The main components of the constant pressure rocket engines are the following: the combustion chamber, also called the firing chamber, and the nozzle located downstream from the combustion chamber (Figure 24). These components can exist in varying quantities, depending on the requirements.
The operating characteristics are as follows: the propellants (fuel and oxidizer) are forced into the combustion chamber in a proper state by means of a sufficient overpressure and are burned there. During the combustion, their chemically bonded energy is converted into heat andin accordance with the related temperature increasealso into a pressure of the combustion gases generated in this manner and enclosed in the combustion chamber. Under the effect of this pressure, the gases of combustion escape out through the nozzle and attain as a result that velocity previously designated as "exhaust velocity." The acceleration of the gas molecules associated with this gain of velocity results, however, in the occurrence of counteracting forces of inertia (counter pressure, similar to pushing away an object!), whose sum now produces the force of "reaction" (Figure 24) that will push the vehicle forward in the same fashion as has already been discussed earlier. The forward thrust is obtained via heat, pressure, acceleration and reaction from the energy chemically bonded in the fuel.
Figure 24. The combustion or firing chamber and the nozzle, the main components of the constant pressure rocket motor.
Key: 1. Escaping gases of combustion; 2. Reactive force; 3. Propellants flowing in, e.g., fuel and oxygen; 4. Combustion chamber.
So that this process is constantly maintained, it must be ensured that continually fresh propellants are injected into the combustion chamber. To this end, it is, however, necessary, as has been stated previously, that the propellant be under a certain overpressure compared to the combustion chamber. If an overpressure is supposed to be available in the tanks, then they would also have to have an appropriate wall thickness, a property, however, that for larger tanks could present problems. Otherwise, pumps will have to be carried on board in order to put the propellants under the required pressure.
Furthermore, related equipment, such as injectors, evaporators and similar units are required so that the on board liquid propellants can also be converted into the state suitable for combustion. Finally, the vehicle designers must also make provisions for sufficient cooling of the combustion chamber and nozzle, for control, etc.
The entire system has many similarities to a constant pressure gas turbine. And similar to that case, the not so simple question also exists in this case of a compatible material capable of withstanding high temperatures and of corresponding cooling options for the combustion chamber and nozzle. On the other hand, the very critical issue of a compressor for a gas turbine is not applicable for the rocket motor.
Similarly, the detonation rocket engine exhibits many similarities to the related type of turbine, the detonation (explosion) gas turbine. As with the latter, the advantage of a simpler propellant injection option must also be paid for in this case by a lower thermal efficiency and a more complicated structure.
Which type of construction should be preferred can only be demonstrated in the future development of the rocket motor. Perhaps, this will also be, in part, a function of the particular special applications of the motor. It would not suffice to have only a motor functioning in completely empty space. We must still have the option of carrying on board into outer space the necessary amounts of energy in the form of propellant. Consequently, we are faced with a critically important question: the construction of the tanks for the fuel and oxidizer.
Key: 1. Following a completed propulsion phase: The rocket is brought to the desired velocity of motion; 2. Remaining "final mass" of the rocket.; 3. Consumed for the propulsion; 4. During the propulsion phase: The rocket is accelerated; 5. Rocket mass (namely, the propellants) is continually expelled.; 6. In the launchready state: The rocket is at rest.; 7. "Initial mass" of the rocket.
How large, in reality, is the amount of propellants carried on board? We know that the propulsion of the rocket vehicle occurs as a result of the fact that it continually expels towards the rear parts of its own mass (in our case, the propellants in a gasified state). After the propulsion system has functioned for a certain time, the initial mass of the vehicle (that is, its total mass in the launchready state) will have been decreased to a certain final mass by the amount of propellants consumed (and exhausted) during this time (Figure 25). This final mass represents, therefore, the total load that was transported by means of the amount of propellants consumed, consisting of the payload, the vehicle itself and the remaining amounts of propellants.
The question is now as follows (Figure 26): How large must the initial mass M0 be when a fixed final mass M is supposed to be accelerated to a velocity of motion v at a constant exhaust velocity c? The rocket equation provides an answer to this question: M0=2.72v/cM.
Key: 1. Velocity of motion; 2. Final mass; 3. Exhaust velocity; 4. Initial mass.
According to the above, the initial mass M0 of a space rocket is calculated as shown below. This mass should be capable of imparting the previously discussed ideal highest climbing velocity of 12,500 meters per second, approximately necessary for attaining complete separation from the Earth.
M0=520 M, for c=2,000 meters per second
M0=64 M, for c=3,000 meters per second
M0=23 M, for c=4,000 meters per second
M0=12 M, for c=5,000 meters per second.
This implies the following: for the case that the exhaust velocity c is, by way of example, 3,000 meters per second, the vehicle, at the beginning of the propulsion phase, must be 64 times as heavy with the propellants necessary for the ascent as after the propellants are consumed. Consequently, the tanks must have a capacity to such an extent that they can hold an amount of propellants weighing 63 times as much as the empty space rocket, including the load to be transported, or expressed differently: an amount of propellants that is 98.5 percent of the total weight of the launchready vehicle.
An amount of propellants of 22 times the weight would also suffice if the exhaust velocity is 4,000 meters per second and only 11 times if the exhaust velocity increases up to 5,000 meters per second. Ninetysix and 92 percent of the total weight of the launchready vehicle is allocated to the propellants in these two cases.
As has been frequently emphasized, the extreme importance of an expulsion (exhaust) velocity as high as possible can clearly be recognized from these values. (The velocity is the expression of the practical energy value of the propellant used!) However, only those rockets that are supposed to be capable of imparting the maximum climbing velocity necessary for the total separation from the Earth must have a propellant capacity as large as that computed above. On the other hand, the "ratio of masses" (ratio of the initial to the final mass of the rocket: M0/M) is considerably more favorable for various types of applications (explained later) in which lower highest velocities also suffice.
In the latter cases from a structural engineering point of view, fundamental difficulties would not be caused by the demands for the propellant capacity of the vehicle and/or of the tanks. By way of example, a space rocket that is supposed to attain the final velocity of v=4,200 meters per second at an exhaust velocity of c=3,000 meters per second would have to have a ratio of masses of M0/M=4, according to the rocket equation. That is, the rocket would have to be capable of storing an amount of propellant that is 75 percent of its total launch weight, a capability that can certainly be achieved from a structural engineering point of view.
To be sure, space rockets of that can carry on board the amounts of propellants necessary for the complete separation from the Earth (according to what has already been stated, the amounts of propellants are 98.5 percent of the launch weight at an exhaust velocity of c=3,000 meters per second), could, for all practical purposes, not be easily realized. Fortunately, there is a trick making it possible to circumvent this structural difficulty in a very simple manner: the socalled staging principle that both Goddard and Oberth recognized independently of one another as a fundamental principle of rocket technology.
In accordance with this principle, the desired final velocity need not be attained with a single rocket; but rather, the space rocket is divided into multiple units (stages), each one always forming the load for the next largest unit. If, for example, a threestage space rocket is used, then it consists of exactly three subrockets: the subrocket 3 is the smallest and carries the actual payload. It forms (including this payload) the load of subrocket 2 and the latter again (including subrocket 3 and its payload) the load of subrocket 1. During ascent, subrocket 1 functions first. As soon as this stage is used up, its empty shell is decoupled and subrocket 2 starts to function. When it is spent, it also remains behind and now subrocket 3 functions until the desired final velocity is attained. Only the latter arrives at the destination with the payload.
Because the final velocities of three subrockets are additive in this process, each individual one must be able to generate only 1/3 of the total required final velocity.
In the case of a 3stage space rocket, which is supposed to attain the highest climbing velocity of 12,500 meters per second necessary for the total separation from the Earth, only a final velocity to be attained of around 4,200 meters per second would consequently be allocated to each subrocket. For that, however, the propellant capacity, certainly implementable from an engineering point of view, of 75 percent (ratio of masses M0/M=4) suffices, as we determined previously, at an exhaust velocity of c=3,000 meters per second, for example. If the individual subrockets can, however, be manufactured, then no doubt exists about the possibility of erecting the complete rocket assembled from all subrockets.
As a precautionary measure, let's examine the absolute values of the rocket masses or rocket weights resulting from the above example. Assume a payload of 10 tons is to be separated from the Earth; the individual subrockets may be built in such a fashion that their empty weight is as large as the load to be transported by them. The weights of the subrockets in tons result then as follows:
Subrocket Load Empty Final weight M Initial weight M0 weight
3 10 10 10 + 10=201) 4 x 20=802)
2 + 3 80 80 80 + 80=160 4 x 160=640
1 + 2 + 3 640 640 640 + 640=1280 4 x 1280=5120
1) The final weight M is equal to the empty weight plus the load when the rocketas in this casefunctions until its propellants are completely consumed.
2) The initial weight M0 is, in this case, equal to 4 times the final weight M because, as has been stated previously in our example, each subrocket approaches the ratio of masses (weights) M0/M=4.
The initial weight of the total space rocket consisting of 3 stages would be 5,120 tons, a number that is not particularly impressive, considering the fact that technology is capable of building, for example, an ocean liner weighing 50,000 tons.
In this fashionby means of the staging principleit would actually be possible to attain any arbitrary final velocity, in theory at least. For all practical purposes in this regard, fixed limitations will, of course, result, in particular when taking the absolute values of the initial weights into consideration. Nevertheless an irrefutable proof is inherent in the staging principle to the effect that it would be fundamentally possible to build space rockets capable of separating from the Earth even with the means available today.
That does not mean the staging principle represents the ideal solution for constructing space rockets in the described form, because it leads to an increase of the dead weight and as a result of the propellants necessary for transportation. This, however, is not now a critical point. Initially, we are only concerned with showing "that it is possible in the first place." Without a doubt every type of space rocket construction, regardless of which one, will have to employ the fundamental concept expressed in the staging principle: during the duration of propulsionfor the purpose of saving propellantsevery part of the vehicle that has become unnecessary must be immediately released (jettisoned) in order not to carry dead weight uselessly and, at the same time, to have to accelerate further with the remaining weight. It is assumed, of course, that we are dealing with space rockets that are supposed to attain greater final velocities.
From a structural engineering point of view, we do not want to conceal the fact that certainly quite a few difficulties will arise as a result of the still significant demands imposed on the capacity of the propellant tanksdespite the staging principle. In this regard, it will be necessary in part to use construction methods deviating fundamentally from the customary ones, because all parts of the vehicle, in particular the tanks, must be made as lightweight as possible. Nevertheless, the tanks must have sufficient strength and stiffness to be able to withstand both the pressure of mass and the atmospheric stagnation pressure during the ascent, taking into account that many of the usual metals become brittle and, therefore, lose strength at the extreme lower temperatures to which the tanks may be exposed.
Moreover in a space ship, a compartment (cell) must exist for housing the pilot and passengers and for storing supplies of the life support necessities and other equipment, as well as for storing freight, scientific devices for observations, etc. The compartment must be airsealed and must have corresponding precautionary measures for artificially supplying air for breathing and for maintaining a bearable temperature. All equipment necessary for controlling the vehicle are also stored in the compartment, such as manual controls for regulating the propulsion system; recorders for time, acceleration, velocity, and path (altitude); and for determining the location, maintaining the desired direction of flight, and similar functions. Even space suits (see the following), hammocks, etc. must be available. Finally, the very important aids for landing, such as parachutes, wings, etc. also belong to the equipment of a space ship.